This invention relates generally to gas turbine engine airfoils, and more particularly to turbine airfoils having reduced tip leakage.
A gas turbine engine includes a compressor that provides pressurized air to a combustor wherein the air is mixed with fuel and ignited for generating hot combustion gases. These gases flow downstream to one or more turbines that extract energy therefrom to power the compressor and provide useful work such as powering an aircraft in flight. In the turbine, an array of airfoil-shaped turbine blades extend radially outwardly from a supporting rotor disk.
The airfoils have opposed pressure and suction sides extending axially between corresponding leading and trailing edges and radially between a root and a tip. The blade tip is spaced closely to a surrounding turbine shroud. The gas pressure difference between the pressure side tip and the suction side tip causes the gas to leak from the pressure side tip through the tip clearance or gap with the shroud, and toward the suction side tip. This tip leakage flow can not produce useful turbine work and will result in performance loss. Thus, maximum efficiency of the engine is obtained by minimizing the tip clearance. However, the degree to which the gap can be reduce is limited by need to allow for differential thermal and mechanical expansion and contraction between the rotor blades and the turbine shroud to prevent undesirable tip rubs.
Accordingly, some prior art turbine blade designs include an offset on the pressure and/or suction sides in order to increase flow resistance through the tip clearance. Examples of such designs are disclosed in U.S. Pat. No. 6,672,829 to Cherry et al., and U.S. Pat. No. 6,790,005 to Lee et al.
Nevertheless, there remains a need for a turbine blade tip which reduces the overall tip leakage flow and thereby increases the efficiency of the turbine.